The present disclosure relates to a gas turbine engine and, more particularly, to a fuel manifold therefor.
Gas turbine engines, such as those that power modern commercial and military aircraft, include a compressor section to pressurize a supply of air, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases and generate thrust.
A fuel manifold system is mounted around a diffuser case of the combustor section and generally includes a multiple of circumferentially distributed fuel injectors that axially project into a combustion chamber to supply fuel thereto. The multiple of circumferentially distributed fuel injectors are connected to multiple fuel supply manifolds that deliver fuel to the fuel injectors though “pigtail” supply assemblies. There is typically one fuel supply manifold for each stage. Thus, each fuel injector may have multiple pigtail supply assembly connections that connect multiple fuel supply manifolds.
In one system, there are two fuel manifold rings, a primary fuel manifold and a secondary fuel manifold. The manifolds are mounted around the circumference of the diffuser case and need to accommodate thermal expansion of the diffuser case. Further, the fuel manifold system typically includes numerous valves, fuel injectors, fuel manifolds, fittings, conduits, pigtail supply assemblies, mounts and other components for both the primary manifold and the secondary manifold that may result in relatively complicated fuel manifold system. Yet, the fuel manifold system must necessarily be contained within an axial and radially constrained packaging space.